1. Field of the Invention
The present invention relates to a shape of a gas passage in an axial-flow gas turbine engine in which a plurality of inlet guide vanes are radially disposed in an annular gas passage defined between an inner peripheral wall and an outer peripheral wail.
2. Description of Background Art
At present, the shapes of the inner peripheral wall and the outer peripheral wall of the gas passage of the inlet guide vanes adopted in the turbines of axial-flow gas turbine engine for an aircraft are mainly cylindrical shapes or conical shapes each formed by a generating line comprising a straight line, in large-sized or medium-sized engines. Also, as a modified shape of the conical shape, there exists a shape formed by an S-shaped generating line having a single inflection point. In small-sized engines, the shape of the outer peripheral wall does not differ from that of large-sized or medium-sized engines, but as for the shapes of the inner peripheral walls, a simple cylindrical shape is generally adopted because the inner peripheral wall is low in design freedom degree since they are small in size.
U.S. Pat. No. 6,283,713 discloses a gas turbine in which the shape of an end wall 33 of a platform 31 constituting the hub of a turbine blade 29 is made to differ at a side of a suction surface 34 and at a side of a pressure surface 35 of the turbine blade 29, thereby reducing the pressure gradient between the adjacent turbine blades 29 to delay the occurrence of vortex and pressure loss.
U.S. Pat. No. 6,669,445 discloses a flow directing assembly in which the surface shape of a platform 16 supporting the inner end in a radial direction of a blade 12 or a vane 12 of a compressor or a turbine of a gas turbine engine is bulged outwardly in the radial direction at a portion in contact with the blade 12 or the vane 12, and is recessed inward in the radial direction at an intermediate portion in the circumferential direction, thereby reducing a shock wave in a transonic region.
U.S. Pat. No. 6,561,761 discloses a compressor flow path in which a flute 40 extending in a gas flow direction is formed in the region sandwiched by adjacent blades 16 in a platform 38 constituting the inner peripheral wall of the blade 16 of the compressor of a gas turbine engine and a shroud 36 constituting the outer peripheral wall, thereby expanding the gas passage to improve efficiency of the compressor.
U.S. Pat. No. 5,466,123 discloses a gas turbine engine in which convex portions and concave portions continuing in the circumferential direction are formed on the inner platform 27 of the nozzle guide vane 20 of a turbine, thereby making uniform distribution in the circumferential direction at the downstream of the nozzle guide vane 20.
It is known that the pressure loss in the inlet guide vane of a turbine of a turbine engine occurs because a secondary flow from a tip toward a hub side occurs due to a pressure difference in the span direction in the suction surface of the inlet guide vane. Thus, the exit flow angle of gas from the trailing edge becomes ununiform in the span direction to reduce the efficiency of the turbine on its rear-stage.